The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling thereof.
In a gas turbine engine, air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases which flow downstream through turbine stages that extract energy therefrom for powering the compressor, and also typically powering a fan for producing propulsion thrust in an aircraft engine application. Each turbine stage includes a stationary turbine nozzle including a row of stator vanes extending radially between outer and inner bands which direct the combustion gases through a downstream row of turbine rotor blades extending radially outwardly from a supporting rotor disk.
The first stage turbine nozzle and blades are subject to the hottest temperature combustion gases discharged from the combustor and require effective cooling for ensuring a suitable useful life thereof. The vanes and blades therefore are hollow for channeling pressurized air bled from the compressor for internal cooling thereof. The vanes and blades typically include rows of inclined film cooling holes through the pressure and suction side surfaces thereof for forming a layer of protective film cooling air to insulate against the hot combustion gases flowing over the vane and blade airfoils.
Since air used in cooling turbine components bypasses the combustor, the overall efficiency of the engine is correspondingly reduced. Accordingly, it is desired to limit the amount of cooling air diverted from the compressor for minimizing the reduction in engine efficiency.
As combustion gas temperature is increased in developing more efficient gas turbine engines, the cooling requirements for the turbines further increase. For example, each turbine blade includes an integral platform at the root thereof which defines a portion of the inner flowpath boundary for the combustion gases. The platforms are typically imperforate and are cooled from their undersides by air channeled in corresponding cavities therebelow.
To further increase platform cooling, the platform may include film cooling holes extending therethrough for film cooling the outer surfaces thereof directly exposed to the hot combustion gases, with the inner surfaces thereof being convection cooled by cooling air circulating within the under platform cavities.
However, film cooling is limited in effectiveness, and the introduction of film cooling holes in the platform of a rotor blade should avoid undesirable stress concentrations which would locally increase stress during operation and correspondingly reduce the useful life of the rotor blades.
Accordingly, it is desired to provide a gas turbine engine turbine blade having improved platform cooling.
A turbine blade includes an integral airfoil, platform, shank, and dovetail, with a pair of holes in tandem extending through the platform and shank in series flow communication with an airflow channel inside the shank. Cooling air discharged through the tandem holes effects multiple, convection, impingement, and film cooling using the same air.